This invention relates to attitude control systems generally, and more particulary to disturbance torque compensation for attitude control systems for stabilizing 3-axis zero-momentum satellites in the absence of attitude reference information from yaw and/or roll inertail references such as gyroscopes.
Modern satellites are widely used for communications and Earth sensing applications. All such satellite applications require control of the satellite's orientation in space so that sensors and antennas may be pointed in appropriate directions. Satellites orbiting the Earth, or other heavenly body (hereinafter Earth), do not maintain a single face directed toward that body without additional control. This control is normally termed attitude control. One type of attitude control includes the use of one or more momentum wheels, which stabilize attitude by providing momentum bias or gyroscopic stiffness. A momentum bias can only indirectly stabilize satellite axes lying in a plane orthogonal to the bias axis, but direct control is not provided. To satisfy the more stringent orientation requirements associated with modern satellite missions, direct control of the satellite axes of rotation is required. Three axes are commonly used: the yaw axis is oriented toward the Earth, the pitch axis is aligned with the satellite's orbit normal, and the roll axis completes the right-handed orthogonal axis set. Those skilled in the art know that other non-orthogonal sets can be used, and that simple transformations relate such sets to the orthogonal set.
The manufacture and launch of satellites is very capital intensive. Consequently, to keep the unit cost of satellite services low, the satellite must be operated for a long time. For this reason, satellite reliability is a major concern requiring such strong measures as redundancy, qualification and pre-launch test.
Typical three axis stabilized satellite attitude control is accomplished by directly sensing the three orthogonal attitudes: Yaw(x), Roll(y), and Pitch(z), and commanding corrective control torques through reaction wheel actuators, or other torque generators. Such control often uses a Earth Sensor Assembly (ESA) to provide roll and pitch attitude information. A gyroscope is used to provide the inertial yaw axis attitude information, because yaw errors (the yaw degree of freedom) are not observable with the ESA. Pitch and roll gyros are often used in ascent and Earth acquisition phases of the satellite's launch. Once the Earth has been acquired, however, the roll and pitch attitude information is derived from the ESA rather than from the gyros.
The ESA assembly can be fabricated without moving parts, and consequently may be very reliable. Even so, redundancy assures availability roll and pitch attitude information throughout the satellite's lifespan. However the yaw gyro is a mechanical device that is prone to failure. Direct redundancy is not normally used because of the gyro's high cost. One additional skew gyro provides the only redundancy for the three orthogonal gyros.
If the yaw gyro fails, redundant skew gyro could be used to provide yaw attitude information, but information derived solely from the skew gyro is contaminated with pitch and roll information. In order to derive yaw attitude information, from the skew gyro, the roll and pitch gyros are enabled to provide corrective signals. If, however, one of the roll or pitch gyros (or both) should also fail, no useful yaw information is available and the satellite attitude becomes uncontrolled. In addition, the satellite pointing error is worsened by environmental disturbance torques.
In low cost satellites in which precision pointing is not required, the cost of the ESA and gyros may be excessive, and other inertial references may not be available. An improved attitude control system is desired.